It is known to provide air from a compressor of an associated gas turbine engine, for the purpose of cooling the engine turbine structure, i.e. at least the respective first high pressure high temperature stage of nozzle guide vanes, and/or the immediately following high pressure high temperature stage of turbine blades. Cooling is achieved by passing air bled from the compressor into passageways formed in the aerofoils of the respective vanes and blades, then ejecting the air into the gas annulus via slots in the trailing edges of the aerofoils.
Some turbine designs incorporate a stage of turbine blades which lie in a gas annulus wherein that part of which surrounding the stage of blades is constructed from a number of segments known as shroud segments. The known method of cooling such a structure is to provide a flow of compressor air over the radially outer surface of the segments, the flow path being defined between those radially outer surfaces and a surrounding casing.